Fluid servo systems are used for many purposes, one being to position the flight control surfaces of an aircraft. In such an application, system redundancy is desired to achieve increased reliability in various modes of operation, such as in a control augmentation or electrical mode.
In conventional electro-hydraulic systems, plural redundant electro-hydraulic valves have been used in conjunction with plural redundant servo valve actuators to assure proper position control of the system's main control servo valve in the event of failure of one of the valves and/or servo actuators, or one of the corresponding hydraulic systems. Typically, the servo actuators operate on opposite ends of a linearly movable valve element of the main control valve and are controlled by the electro-hydraulic valves located elsewhere in the system housing. Although the servo valve actuators, alone or together, advantageously are capable of driving the linear movable valve element against high reaction forces, such added redundancy results in a complex system with many additional electrical and hydraulic elements necessary to perform the various sensing, equalization, timing and other control functions. This gives rise to reduced overall reliability, increased package size and cost, and imposes added requirements on the associated electronics.
An alternative approach to the electro-hydraulic control system is an electro-mechanical control system wherein a force motor is coupled directly and mechanically to the main control servo valve. In this system, redundancy has been accomplished by mechanical summation of forces directly within the multiple coil force motor as opposed to the conventional electro-hydraulic system where redundancy is achieved by hydraulic force summing using multiple electro-hydraulic valves and actuators. If one coil or its associated electronics should fail, its counterpart channel will maintain control while the failed channel is uncoupled and made passive. Such alternative approach, however, has a practical limitation in that direct drive forces motors utilizing state of the art rare earth magnet materials are not capable of producing desired high output forces at the main control servo valve within acceptable size and weight limitations.
In aircraft flight control systems, it also is advantageous and desirable to provide for controlled recentering of the main control servo valve in the event of a total failure or shut-down of the electrical operational mode. This is particularly desirable in those control systems wherein a manual input to the main servo valve is provided in the event that a mechanical reversion is necessary after multiple failures have rendered the electrical mode inoperative. In known servo systems of this type, the manual input may operate upon the spool of the main servo valve whereas the electrical input operates upon the movable sleeve of the main servo valve.
Upon rendering the electrical mode inactive, it is necessary to move the valve sleeve to a neutral or centered position and lock it against movement relative to the valve spool controlled by the manual input. Heretofore, this has been done by using a centering spring device which moves the valve sleeve to its centered or neutral position and a spring biased plunger that engages a slot in the valve sleeve to lock the latter against movement. The plunger normally is maintained out of engagement with the slot during operation in the electrical mode by hydraulic system pressure, and may have a tapered nose that engages a similarly tapered slot in the valve sleeve to assisat in centering the valve sleeve.